Nowadays gas turbines are used as a power source of many kinds of machinery and equipment. For example, they are used for power plant applications by connecting a generator to their main shaft, or used as engines utilizing the gas turbines as a power source of transportation such as airplanes and the like.
FIG. 18 is a schematic drawing of a gas turbine engine. A gas turbine engine GT shown in FIG. 18 has an intake fan Kf installed to an air intake port; a compressor Cp compressing the intake air; a combustor Bs burning fuels by using the air compressed by the compressor Cp; a turbine Tb driven by combustion gas ejected from the combustor Bs; and a nozzle Nz ejecting combustion gas to generate thrust. The intake fan Kf, the compressor Cp and the turbine Tb are connected to each other by the same rotating shaft. Driving the turbine Tb with combustion gas turns the rotating shaft, driving the intake fan Kf and the compressor Cp.
The turbine Tb has stationary vanes and rotating blades. The stationary vanes rectify flow of the combustion gas blown thereto, while the rotating blades rotate with the rotating shaft at their center by having the combustion gas blow thereto. Combustion gas blowing to the stationary vanes and the rotating blades is very high temperature gas, thus generating nonconformances such as thermal deformation, damage and the like due to heat. In order to prevent these nonconformances from occurring, the blades are cooled.
A method of cooling the blades is disclosed in which a part of the compressed air, which is compressed by the compressor Cp and supplied to the combustor Bs, is supplied to the turbine Tb to be utilized as a refrigerant for cooling of the blades. In this method, the blades are cooled by having the cooling air flow inside the blades. Among methods of cooling the blades are a film cooling method, an impingement cooling method, a transpiration cooling method and the like.
In order to cool turbine blades of high temperature gas turbine engines, after implementation of the impingement cooling method, the film cooling method is also used. FIG. 19 shows a cross-sectional view of a cooled blade of a high temperature gas turbine engine. The cooled turbine blade B shown in FIG. 19 has a cooling passageway 91 formed inside thereof where the cooling air flows. Inside the cooling passageway 91 is mounted an insert 92, having a multiple number of small holes for ejecting the cooling air, leaving a space apart from an inner wall surface 901 of a blade wall 90. The cooling air flowing into the cooling passageway 91 is blown approximately uniformly to the inner wall surface 901 of the cooling passageway 91 through small holes 921 in the insert 92 and cools the inner wall surface 901 and the neighborhood of the inner wall, which is so-called an impingement cooling
Additionally, the cooled blade B has film-cooling holes 93 formed therein, penetrating from the cooling passageway 91 to the outside. As a result, the cooling air which completes the impingement cooling will be ejected to an external wall surface 902 of the cooled turbine blade B through the film-cooling holes 93, forming a cooling seal and cooling the cooled turbine blade B from the outside.
The above-mentioned cooling methods are widely adopted to gas turbines for industrial use, such as gas turbines used in power plant applications and the like.
Table 1 shows a comparison of various kinds of parameters of cooled blades of a small type of gas turbine engine used for airplanes and the like with those of cooled blades of gas turbines for industrial use such as power generation and the like. Parameters shown in Table 1 include hole diameter d of film-cooling holes, chord length C, blade wall thickness 6 (See FIG. 19 and FIG. 21 described later.), ratio d/C of the film hole diameter versus the chord length and ratio d/δ of the film-cooling hole diameter versus the blade wall thickness. FIG. 20 shows a cross-sectional view of a blade and a graph showing a distribution of static pressure around the blade. The distribution of static pressure around a blade as shown in the graph is for a blade of no dimensions.
TABLE 1Small Type of GasGas TurbinesTurbine Enginesfor Industrial UseFilm Hole Diameter0.40.8d (mm)Chord Length30200C (mm)Blade Wall Thicknessδ (mm)1.25.0d/C0.0130.004d/δ0.330.16
As shown in Table 1, in gas turbines for industrial use, the ratio d/C of the film-cooling hole diameter versus the chord length is 0.004, while in small type of gas turbine engines, the ratio is 0.013, which is a large value. In the graph in FIG. 8, blades are of no dimensions, and the ratio of film-cooling holes of small gas turbine engines and the ratio of those of gas turbines for industrial use are depicted as 81 and 82 in the graph. From the graph, it can be seen that the ratio 81 of the film-cooling holes of small gas turbine engines versus blades is larger than the ratio 82 of those of gas turbines for industrial use versus blades, and it also can be seen that pressure fluctuation at the outlet of the film-cooling holes of small gas turbine engines becomes higher. As a result, combustion gas becomes easier to flow reversely through the film-cooling holes. Therefore, in order to prevent a back flow, it is necessary to increase a margin for the back flow. Namely, it is necessary to make a large amount of the cooling air flow.
FIG. 21 shows a cross-sectional view of a film-cooling hole. A film-cooling hole 93 is formed so as to incline to the blade wall 90 at a predetermined angle. In the neighborhood of an inner wall surface 901 of a curved surface 931 on the inner wall side of the film-cooling hole 93, exfoliation of the cooling air occurs, wherein the cooling air flows away from the curved surface 931 on the inner wall side of the film-cooling hole 93, and subsequently, flow of the cooling air re-adheres to the curved surface 931 on the inner wall side of the film-cooling hole 93 in the proximity of the external wall surface 902 of the film-cooling hole 93.
When the film-cooling hole 93 does not have a sufficient length longer than a predetermined length, the flow of the cooling air does not re-adhere to the film-cooling hole 93, and high temperature working fluid flows reversely from the outside to a portion where the exfoliation of the cooling air flow occurs. This diminishes cooling performance.
Gas turbines for industrial use have a ratio d/δ of the film-cooling hole diameter versus the blade wall thickness which is 0.16, while small gas turbine engines has a ratio of 0.33, which is a large value. Therefore, the film-cooling hole 93 has a difficulty in having a sufficient length for the cooling air flow to re-adhere thereto, and there is a high potentiality of occurrence of a back flow, thus giving an adverse effect on cooling performance of the film cooling. Additionally, it is difficult to apply shaped cooling holes.
Also, as shown in FIG. 20, static pressure around the blade becomes lower toward the downstream side of the flow around the blade. As a result, the working fluid flowing into the inside of the cooled turbine blade B flows toward the downstream side through a gap between the insert 92 and the inner wall 901 of the cooling passageway 91, thereby significantly deteriorating cooling performance of the impingement cooling.